The fundamental physical principle which all space vehicles utilize for propulsion is based on Newton's third law of motion: "For every action, there is an equal and opposite reaction." Thus, by expelling reaction mass (i.e., propellent) at high velocity u through an exhaust nozzle in one direction, a propulsive reaction thrust T is generated in the opposite direction. The magnitude of this thrust is given by EQU T=mu (1)
where m denotes the mass flow rate at which the propellent leaves the exhaust nozzle. This is the fundamental principle which all prior art space propulsion systems use for generating accelerating or decelerating propulsive thrust. (Decelerating retro thrust is obtained by turning the vehicle around and expelling the propellent in the direction of motion.)
If M.sub.1 and M.sub.2 denote the total mass of a space vehicle before and after a propulsive maneuver to achieve a velocity change .DELTA.V, the "mass ratio" M.sub.1 /M.sub.2 is related to .DELTA.V by the well known "rocket equation" given by ##EQU1##
Equations (1) and (2) are among the most important in astronautics because they essentially determine the capabilities and limitations of all space vehicles propelled by prior art propulsion systems. For example, since the total amount of propellent mass m.sub.p that must be carried on-board a space vehicle in order to achieve a total velocity change .DELTA.V is given by M.sub.1 -M.sub.2, it follows from equation (2) that EQU m.sub.p =M.sub.2 [exp(.DELTA.V/u)-1] (3)
Thus, the efficiency of a vehicle's propulsion system increases exponentially with increasing exhaust velocity.
Unfortunately, it is impossible to construct a chemical rocket engine that can produce an exhaust velocity greater than approximately 4.70 km/sec (which corresponds to a specific impulse of 480 sec). This limit is imposed by basic laws of thermodynamics that cannot be violated.
This limit on exhaust velocity has a profound effect on all chemical rockets and it is the principle reason why an Earth-to-orbit launch vehicle requires so much more propellent mass than payload mass. For example, in order to place a payload into a 200 km high circular low Earth orbit (LEO), a launch vehicle has to generate a total .DELTA.V of about 13 km/sec (when atmospheric drag and gravity losses are taken into consideration). Hence, in view of equation (2), the lowest possible mass ratio for a single stage reusable chemical launch vehicle will be about 16.0. If the vehicle's dry mass is 150,000 kg and its payload mass is 30,000 kg, the required pre-launch propellent mass will be 2,880,000 kg. Thus, the payload mass is only 1.01% of the propellent mass.
Since the demands of reaching LEO by chemical rocket propulsion are so high, a reusable, single stage launch vehicle will not be capable of carrying out any significant amount of additional propulsive maneuvers after reaching LEO. Any payload requiring insertion into a higher energy orbit will be taken out of the launch vehicle in LEO and placed aboard a specially designed, reusable, space-based orbital transfer vehicle (OTV). The OTV will then take the payload from LEO and deliver it to the higher energy orbit. In the meantime, the launch vehicle is returned back to Earth where it is refueled and reloaded with another payload.
Most of the required higher energy orbits will be geosynchronous Earth orbit (GEO). If the thrust to weight ratio (T/W) of the loaded OTV is relatively high (e.g.,&gt;0.1 which is required for short flight times) the minimum .DELTA.V required to reach GEO from a 200 km high LEO with an inclination of 28.6.degree. is about 4.30 km/sec. Thus, a round trip mission from LEO to GEO and back to LEO will require a minimum .DELTA.V of about 8.60 km/sec. If the drymass of a chemically propelled OTV is 30,000 kg, the initial propellent mass m.sub.p required to transport a 30,000 kg payload from LEO to GEO and return back to LEO with no payload (and empty fuel tanks) will be 202,000 kg. (In view of the required propellent mass, an OTV dry mass of 30,000 kg is very conservative because the mass of empty cryogenic storage tanks is usually 12% of the propellent mass that must be stored in them.) Since all of the propellent needed to operate the OTV must be transported up from the Earth's surface by ground-to-orbit launch vehicles, the total mass that must be transported up from the Earth's surface in order to place a 30,000 kg payload in GEO will be 232,000 kg. Thus, only 12.9% of the total mass brought up to LEO by the ground-to-orbit launch vehicles will be actual hardware payload. The remaining 87.1% of the mass will be propellent for the OTV. Since it was demonstrated that transporting any payload up to LEO from the Earth's surface is a very difficult and costly task and requires the expenditure of enormous quantities of propellent, the fact that about 87% of the total payload mass delivered to LEO must be wasted on propellent for the OTV makes the prospects for large scale commercial space travel beyond LEO using prior art propulsion systems very discouraging.
The above calculations show that in order to deliver payloads to GEO from LEO by chemically propelled reusable, space-based OTVs, almost all of the total mass that must be transported to LEO from the Earth's surface will be propellents for the OTVs. In fact, for missions to GEO by chemical OTVs, about 90% of the total weight lifting capability of the ground-to-orbit launch vehicle will be wasted by having to carry up the propellents for the OTVs. Only about 10% of the launch vehicle's total payload capability can be allocated for the actual hardware. This unfortunate situation is inherent in all chemical propulsion systems and is not likely to improve with future developments. Although electric propulsion systems do offer a substantial reduction in propellent requirements, they suffer from having inherently long flight times. Thus, for short flight time, high mass missions to GEO (such as manned flights) it appears that chemically propelled OTVs are necessary and unavoidable. This will inevitably result in burdening the ground-to-orbit launch vehicle with having to carry enormous quantities of liquid rocket propellents (for the OTVs) because the fuel-to-hardware mass ratio of the launch payload will be about 10 to 1.
For detailed technical information on prior art OTV studies, the reader is referred to the following references: (1) "Propulsion Options for Space-Based Orbital Transfer Vehicles," Journal of Spacecraft & Rockets, Vol. 16, No. 5, Sept-Oct. 1979 by J. J. Rehder et al; (2) "Orbit Transfer Propulsion and Large Space Systems," Journal of Spacecraft & Rockets, Vol. 17, No. 6, Nov-Dec. 1980, by K. E. Kunz; (3) "Preliminary Design for a Space-Based Orbital Transfer Vehicle," Journal of Spacecraft & Rockets, Vol. 17, No 3, May-June 1980, by I. O. Mac Conachie and J. J. Rehder; and (4) "Electric vs Chemical Propulsion for a Large-Cargo Orbit Transfer Vehicle," Journal of Spacecraft and Rockets, Vo. 16, No. 3, May-June 1979, by J. J. Rehder and K. E. Wurster.
Unfortunately, chemically propelled OTVs will essentially rule out the possibility of space flights to GEO by ordinary private citizens because they will be too expensive. For example, the minimum cost for transporting a 100 kg passenger from the Earth's surface to an orbiting geosynchronous space station via a chemical OTV will be about $1,000,000 (in 1980 dollars). There are not many private citizens who would be able or willing to pay this amount of money for a one-way trip to GEO. Since commercial flights to the Moon or into interplanetary space will be much more expensive than trips to GEO, they will be economically impossible for ordinary private citizens. Thus, for all practical purposes, large scale commercial space travel beyond LEO will be an economic impossibility using prior-art propulsion systems.
In an attempt to circumvent the fundamental problem of having to transport huge quantities of chemical propellents up from the Earth's surface to support chemical OTVs, Gerard O'Neill proposed constructing a base on the Moon, extracting oxygen from Moon rocks, and transporting it to the OTVs from the Moon's surface instead. (See, "The Colonization of Space," Physics Today, Sept. 1974, pp. 32-40.) According to this plan, only the hydrogen component of the propellents would have to be transported up from the Earth's surface. Consequently, since the oxygen-to-hydrogen mass ratio for chemical O.sub.2 /H.sub.2 rocket engines is 8 to 1, the process of having to bring up only the hydrogen component for the OTVs, reduces the total propellent mass that would otherwise have to be transported up from the Earth's surface by a factor of 8. Unfortunately, the problems of maintaining a Moon base, extracting oxygen from Moon rocks, and transporting it from the Moon's surface to the OTVs, introduces operational complexities that turn out to be very costly. (See, "Should We Make Products On The Moon," Astronautics & Aeronautics, June 1983, pp. 80-85 by W. F. Carroll et al). This proposal for solving the propellent refueling problem of chemically propelled OTVs by extracting oxygen from Moon rocks demonstrates the magnitude of the problem and the extraordinary lengths engineers are willing to go in an attempt to circumvent it.
The solution to the refueling problem of OTVs proposed herein by the applicant is radically different from all others in the prior art in that it is based on setting forth a fundamentally new propulsion concept for propelling any reusable space-based vehicle--whether it is an interorbital transfer vehicle moving around the Earth or an interplanetary transfer vehicle moving around the solar system. This propulsion concept is based upon utilizing the atmosphere of a planet as propulsive working fluid and designing the propulsion system to be self-refueling. This is accomplished by first recognizing three simple facts: (1) a reusable space-based vehicle requires just as many decelerating retro propulsive maneuvers as accelerating propulsive maneuvers; (2) that any propulsive maneuver will be more efficient if applied when the vehicle is moving closest to the central body when its velocity is maximum; and (3) that most of the required decelerating retro maneuvers can be generated by a reverse application of the fundamental theory of rocket propulsion by dipping into and ingesting atmospheric gas instead of expelling gas. The third fact represents the most important and the most far-reaching observation. Thus, instead of generating decelerating retro thrust by the usual method of expelling gas through a rocket nozzle pointed in the direction of motion, the proposed system ingests gas through a diffuser intake nozzle pointed in the direction of motion while the vehicle passes through the upper regions of a planet's atmosphere. The vehicle thereby transfers a large portion of its initial momentum to the collected gas which was initially at rest in the upper atmosphere This transfer of momentum provides the vehicle with a free decelerating retro thrust by a reverse application of the theory of classical rocket propulsion.
By designing the intake diffuser nozzle to have a maximum diameter greater than the vehicle's diameter, all aerodynamic drag effects are essentially eliminated and the magnitude of the retro thrust can be calculated by a reverse application of equation (1). In this case the propellent mass flow rate m refers to the intake mass flow rate m.sub.in of atmospheric gas scooped up and ingested by the vehicle, and the flow velocity u refers to the gas inlet velocity u.sub.in. The gas inlet velocity u.sub.in is equal to the vehicle's instantaneous velocity relative to the passing planet. It can be approximated by the vehicle's perigee velocity (when its altitude is minimum)
The vehicle's velocity change .DELTA.V=V.sub.1 -V.sub.2 can be calculated by applying the principle of conservation of momentum. Hence, if M.sub.1, V.sub.1 and M.sub.2, V.sub.2 denote the vehicle's total mass and velocity before and after scooping up an amount of gas m.sub.p =M.sub.2 -M.sub.1 respectively, then M.sub.1 V.sub.1 =M.sub.2 V.sub.2. Consequently, the mass ratio is given by ##EQU2## Notice that in this case, the mass ratio M.sub.1 /M.sub.2 is always less than 1 (because M.sub.2 &gt;M.sub.1) and that in order for V.sub.2 .fwdarw.O, M.sub.2 .fwdarw..infin.. Equation (4) represents the analogue of the basic rocket equation (2) where M.sub.2 &lt;M.sub.1 (since, in this equation, the propellent mass is expelled instead of being ingested.) In deriving equation (4), the gravitational effects are omitted for simplification.
The retro velocity change .DELTA.V=V.sub.1 -V.sub.2 is given by ##EQU3## and the amount of propellant gas m.sub.p ingested during the gas scooping retro maneuver is given by ##EQU4## Equation (6) is the analogue of equation (3).
The atmospheric gas enters the vehicle through the intake diffuser nozzle, cooled, liquefied and stored cryogenically on-board the vehicle. The vehicle uses the collected atmospheric gas as its propulsive working fluid to generate accelerating forward thrust by expelling the gas at high velocity u.sub.out via the usual application of the theory of classical rocket propulsion. It follows directly from equations (3) and (6) that if the exhaust velocity u.sub.out can be made to exceed the maximum inlet velocity u.sub.in =V.sub.1, the vehicle will become self-refueling. Consequently, the gas expended during the vehicle's forward accelerating maneuvers is automatically replenished by the incoming gas during the vehicle's decelerating retro maneuvers.
The crucial operating condition u.sub.out &gt;u.sub.in can be satisfied by employing a high power electromagnetic propulsion system. Several different types can be used and are generally known as high power plasma accelerators. The most popular of these are referred to as magnetohydrodynamic (MHD) accelerators and magnetoplasmadynamic (MPD) accelerators. However, there is one relatively unknown class of plasma accelerators that is ideally suited for the proposed self-refueling propulsion system. They are called electron cyclotron resonance (ECR) accelerators. Although these accelerators are powered by microwaves, they can be used to accelerate any ionized gas to very high exhaust velocities with very high power densities. Thus, it is possible to design a high power ECR accelerator that can be used to accelerate a multitude of different gases. Although some gases will be more efficient to use than others because of their different ionization potentials, all of the gases found on all of the planets can be accelerated by a suitably designed ECR accelerator. Since nitrogen has a relatively low ionization potential, it will be an ideal working fluid for an ECR accelerator system. Thus, since nitrogen represents the major portion of the Earth's atmosphere (76.7%), a self-refueling propulsion system will operate very efficiently for Earth-based reusable OTVs.
The application of the proposed self-refueling propulsion system (that is based on a generalized theory of classical rocket propulsion) is ideally suited for large manned reusable space-based OTVs orbiting the Earth, as well as for manned reusable interplanetary vehicles moving from planet to planet. Since no rocket propellents have to be transported up from the Earth's surface (or from the surface of any other celestial body) to refuel these vehicles, the concept represents a revolutionary propulsion breakthrough in achieving economical commercial space travel throughout the entire solar system.